U.S. patent number 7,597,838 [Application Number 11/027,403] was granted by the patent office on 2009-10-06 for functionally gradient sic/sic ceramic matrix composites with tailored properties for turbine engine applications.
This patent grant is currently assigned to General Electric Company. Invention is credited to Douglas Melton Carper, Toby George Darkins, Jr., James Dale Steibel, Suresh Subramanian.
United States Patent |
7,597,838 |
Subramanian , et
al. |
October 6, 2009 |
**Please see images for:
( Certificate of Correction ) ** |
Functionally gradient SiC/SiC ceramic matrix composites with
tailored properties for turbine engine applications
Abstract
A ceramic matrix composite with a ceramic matrix and a gradient
layering of coating on ceramic fibers. The coating typically
improves the performance of the composite in one direction while
degrading it in another direction. For a SiC-SiC ceramic matrix
composite, a BN coating is layered in a gradient fashion or in a
step-wise fashion in different regions of the article comprising
the ceramic. The BN coating thickness is applied over the ceramic
fibers to produce varying desired physical properties by varying
the coating thickness within differing regions of the composite,
thereby tailoring the strength of the composite in the different
regions. The coating may be applied as a single layer as a
multi-layer coating to enhance the performance of the coating as
the ceramic matrix is formed or infiltrated from precursor
materials into a preform of the ceramic fibers.
Inventors: |
Subramanian; Suresh (Mason,
OH), Steibel; James Dale (Mason, OH), Carper; Douglas
Melton (Trenton, OH), Darkins, Jr.; Toby George
(Loveland, OH) |
Assignee: |
General Electric Company
(Schenectady, NY)
|
Family
ID: |
36640786 |
Appl.
No.: |
11/027,403 |
Filed: |
December 30, 2004 |
Prior Publication Data
|
|
|
|
Document
Identifier |
Publication Date |
|
US 20060147688 A1 |
Jul 6, 2006 |
|
Current U.S.
Class: |
264/640;
428/297.1; 428/294.1; 428/293.4; 264/642; 264/605; 501/95.2;
428/297.7; 428/297.4 |
Current CPC
Class: |
F01D
5/282 (20130101); C04B 35/565 (20130101); F01D
5/284 (20130101); C04B 35/62871 (20130101); C04B
35/62884 (20130101); C04B 35/62873 (20130101); C04B
35/80 (20130101); C04B 35/62868 (20130101); C04B
35/62897 (20130101); C04B 35/62894 (20130101); C04B
35/62865 (20130101); C04B 35/62863 (20130101); Y10T
428/24993 (20150401); Y10T 428/24975 (20150115); F05D
2300/2261 (20130101); Y10T 428/24994 (20150401); Y10T
428/265 (20150115); Y10T 442/2975 (20150401); Y10T
428/249931 (20150401); C04B 2235/5264 (20130101); Y10T
428/249929 (20150401); C04B 2235/524 (20130101); C04B
2235/5248 (20130101); F05D 2300/2283 (20130101); Y02T
50/60 (20130101); C04B 2235/5244 (20130101); C04B
2235/5224 (20130101); F05D 2300/603 (20130101); Y10T
428/249928 (20150401); F05C 2203/0839 (20130101); Y10T
428/249939 (20150401); Y10T 428/249924 (20150401); C04B
2235/80 (20130101); Y10T 428/2495 (20150115); C04B
2235/522 (20130101); Y10T 428/249941 (20150401); C04B
2235/5232 (20130101) |
Current International
Class: |
B32B
17/12 (20060101); B32B 15/04 (20060101); C04B
33/34 (20060101); C04B 35/00 (20060101); C04B
35/64 (20060101) |
Field of
Search: |
;428/293.4,293.7,297.4,297.7 ;264/605,642,640 ;501/95.2 |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Tarazano; D. Larazano
Assistant Examiner: Thompson; Camie S
Attorney, Agent or Firm: McNees Wallace & Nurick,
LLC
Government Interests
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT
This invention was made with Government support under Contract No.
N00421-00-3-0536. The government may have certain rights to the
invention.
Claims
What is claimed is:
1. A method of manufacturing a ceramic fiber preform for use in a
ceramic matrix composite, the method comprising: providing a fiber
preform having a plurality of ceramic fibers; defining a plurality
of regions of the fiber preform on the basis of variations in the
desired level of resistance to in-plane stress and variations in
the desired level of resistance to interlaminar stress among
corresponding regions of the ceramic matrix composite, the
plurality of regions including a first region and a second region;
applying BN to the first region, yielding a first BN layer having a
thickness in the range of about 0.3 micron to about 1.0 micron;
applying BN to the second region, yielding a second BN layer having
a thickness of less than or equal to about 0.3 micron; applying
Si-doped BN to the first BN layer, yielding a first Si-doped BN
layer having a thickness in the range of about 0.4 micron to about
0.6 micron; applying Si-doped BN to the second BN layer, yielding a
second Si-doped BN layer having a thickness of less than or equal
to about 0.3 micron; applying SiN to the first Si-doped BN layer,
yielding a first SiN layer having a thickness in the range of about
0.75 micron to about 1.25 micron, thereby forming a first
multilayer fiber coating corresponding to the first region; and
applying SiN to the second Si-doped BN layer, yielding a second SiN
layer having a thickness of about 0.5 micron, thereby forming a
second multilayer fiber coating corresponding to the second
region.
2. The method of claim 1, wherein the ceramic fiber is selected
from the group consisting of SiC, Al.sub.2O.sub.3 and Si--N--C.
3. The method of claim 1, wherein the thickness of the first BN
layer is about 0.3 micron.
4. The method of claim 1, wherein the thickness of the second BN
layer is about 0.3 micron.
5. The method of claim 1, wherein the thickness of the second
Si-doped BN layer is about 0.3 micron.
6. The method of claim 1, further including applying carbon to the
first SiN layer, yielding a carbon layer having a thickness of less
than or equal to about 0.1 micron.
7. The method of claim 1, further including applying carbon to the
first SiN layer, yielding a carbon layer having a thickness of
about 0.05 micron.
8. The method of claim 1, wherein the plurality of regions includes
a third region, and further including applying BN to the third
region, yielding a BN monolayer having a thickness of about 1.0
micron.
9. The method of claim 1, wherein the plurality of ceramic fibers
essentially includes only ceramic fibers having a diameter in the
range of about 5 microns to about 20 microns.
10. The method of claim 1, wherein the plurality of ceramic fibers
essentially includes only ceramic fibers having a diameter in the
range of about 10 microns to about 15 microns.
11. The method of claim 10, wherein the plurality of ceramic fibers
consists essentially of ceramic fibers composed of SiC.
12. The method of claim 1, wherein the first region and the second
region are contiguous.
13. The method of claim 12, wherein a gradient is defined by the
first multilayer fiber coating and the second multilayer fiber
coating.
14. The method of claim 1, wherein the first region and the second
region are not contiguous.
Description
FIELD OF THE INVENTION
The present invention relates generally to ceramic matrix composite
materials . More particularly, this invention is directed to the
application of a gradient layer of boron nitride (BN) applied to
silicon carbide (SiC) fibers within a fiber reinforced ceramic
matrix composite to improve interlaminar strength.
BACKGROUND OF THE INVENTION
A gas turbine engine includes a compressor, a combustor and a
turbine. The compressor includes a plurality of disks, each with
arcuate blades extending from its periphery. The disks rotate
rapidly on a shaft, drawing in air, compressing it and moving the
highly compressed air downstream toward the combustor. In the
combustor, the compressed air is mixed with metered fuel which is
burned, generating hot gases. The hot gases flow to the turbine
which comprises at least one disk having arcuate blades extending
from its periphery. Energy is extracted from the hot gases by the
blades, the hot gases striking the blades causing the disks to
turn, which in turn rotates the shaft, powering the engine. The
remaining gases passing through the turbine generate thrust to
propel an aircraft.
The materials used in the turbine section, because of their
exposure to high temperatures and because of the rapid rotation,
have typically been comprised of high temperature superalloys.
However, because of their light weight and high temperature
capabilities, ceramic composite materials, such as SiC/SiC ceramic
matrix composites, which exhibit favorable characteristics, have
been considered for use in the turbine portion of the engine, such
as in turbine engine blade applications. One of the drawbacks of
this material has been its poor interlaminar properties. The
primary cause for low interlaminar strength is the presence of a
boron nitride coating, which is typically applied over the fibers
to form an interface between the fibers and matrix, thereby
increasing fracture toughness by allowing the load to be
transferred to the fibers and absorb energy by promoting crack
propagation along the fibers, or within the weaker fiber coating.
This low interlaminar strength, by improving fracture toughness,
reduces the tendency of the material to suddenly fail in a brittle
mode. In many of the hot section applications such as combustion
liners, HPT vanes, LP blades and shrouds, the thermal gradients and
mechanical loads can result in significant local interlaminar
stresses. Therefore, it is desirable to enhance the interlaminar
strength of ceramic composites in local areas for many of these
applications.
A number of techniques have been used in the past to manufacture
turbine engine components, such as turbine blades using ceramic
matrix composites. However, such turbine components, under normal
operating conditions, are not subjected to uniform stress patterns,
instead experiencing varying degrees of local stresses at different
times and at different locations within the part during normal
turbine operation. A turbine blade generally has a dovetail
portion, an airfoil portion opposite the dovetail portion and an
optional platform located between the dovetail portion and the
airfoil portion. In the dovetail portion of turbine blades,
relatively higher tensile stress regions are located in the
outermost portion of the dovetail section. Ideally, the CMC
component should be designed such that the component has a higher
tensile strength in the region experiencing the higher tensile
stresses. One method of manufacturing CMC components, set forth in
U.S. Pat. Nos. 5,015,540; 5,330,854; and 5,336,350; incorporated
herein by reference and assigned to the assignee of the present
invention, relates to the production of silicon carbide matrix
composites containing fibrous material that is infiltrated with
molten silicon, the process herein referred to as the Silcomp
process. The fibers generally have diameters of about 140
micrometers or greater, which prevents the manufacture of
intricate, complex shapes, such as turbine blade components, by the
Silcomp process.
Another technique of manufacturing CMC turbine blades is the method
known as the slurry cast melt infiltration (MI) process. A
technical description of such a MI method is described in detail in
U.S. Pat. No. 6,280,550 B1, which is assigned to the Assignee of
the present invention and which is incorporated herein by
reference. In one method of manufacturing using the MI method, CMCs
are produced by initially providing plies of balanced
two-dimensional (2D) woven cloth comprising silicon carbide
(SiC)-containing fibers, having two weave directions at
substantially 90.degree. angles to each other, with substantially
the same number of fibers running in both directions of the weave.
By "silicon carbide-containing fiber" is meant a fiber having a
composition that includes silicon carbide, and preferably is
substantially only silicon carbide. The fiber may have a silicon
carbide core surrounded with carbon, or in the reverse, the fiber
may have a carbon core surrounded by, or encapsulated with, silicon
carbide. These examples are exemplary of the term "silicon
carbide-containing fiber" and are not limited to this specific
combination. Other fiber compositions are contemplated, so long as
they include silicon carbide.
Prior ceramic matrix composites, such as U.S. Pat. No. 4,642,271,
to Rice may be suitable for producing a homogenous composite with
favorable toughness characteristics and other inplane properties,
but lack the interlaminar properties required for many turbine
engine applications. Typical methods of improving interlaminar
strength of SiC/SiC composites have utilized through thickness
fiber reinforcement. T-forming and Z-pinning are examples of
techniques used to introduce load carrying fibers in the
through-thickness directions of composites and, thus, enhance
interlaminar strength within desired regions. T-forming, as set
forth in U.S. Pat. No. 6,103,337 to Moody, is a method by which
fibers are inserted directly into a preform so that spacing, depth
of penetration, and orientation can be controlled to produce 3-D
fiber architectures with improved interlaminar strength. Z-pinning
is a technique used to reinforce a composite structure to prevent
various layers rigidly connected to one another and from
delaminating. These methods, however, require trade-offs in
in-plane mechanical properties and result in significant increases
in fiber and/or manufacturing costs.
Accordingly, there is a need for a method of producing a composite
that possesses regions of favorable in-plane properties and regions
of favorable interlaminar properties, thereby overcoming the
inadequacies of the prior art.
SUMMARY OF THE INVENTION
The present invention is directed to a ceramic matrix composite
with improved interlaminar strength in selected regions. The
present invention provides a method of applying an interface
coating thickness to a fiber preform. The physical properties of a
ceramic composite material manufactured from the preform of the
present invention will vary along or across a composite section. In
this manner, desired physical properties of the material can be
varied in different regions of the composite article to correspond
to, for example, the actual stresses experienced in the specific
regions of the article.
In one embodiment, the present invention provides a method of
manufacturing a ceramic fiber preform, for use in a ceramic matrix
composite, that includes forming a fiber preform with a plurality
of ceramic fibers. The ceramic fibers are provided with an
interface coating, such as BN. The thickness of the interface
coating applied to the ceramic fiber is dependent upon the location
of the fiber within the preform. The coating thickness applied to
the fibers is variable. The thickness of the coating on the fiber
is related to the physical properties required in the region of the
preform in which the fiber is located. In this manner, the coating
applied to the ceramic fibers is graded, that is to say, the
coating applied to the fibers is not uniform across the preform,
but rather varies in thickness depending upon the region of the
preform in which the fiber is assembled, thereby producing an
article with non-uniform, or graded, mechanical properties through
its section.
In another embodiment, the present invention provides a method of
manufacturing a component or article, such as an article for use in
a gas turbine engine, the component having tailored mechanical
properties. The method comprises first identifying distinct regions
of a composite structure wherein different mechanical properties
are required, and predicting the required mechanical properties.
The method then requires a determination of a desired coating
thickness applied to ceramic fibers to achieve the predetermined
mechanical properties, with different thicknesses applied to
achieve different mechanical properties. Fibers having the required
coating thickness to achieve the predetermined mechanical
properties are assembled within the preform so as to have regions
in which coating thickness varies to produce different mechanical
properties. The preform is then formed into a composite structure
by any convenient method to produce a composite component having
graded mechanical properties, the properties varying across the
structure as a result or the varying coating applied to the ceramic
fibers. The preform and the finished article may have graded
mechanical properties that are continuously graded, in that the
mechanical properties vary incrementally across the article due to
slight variations in coating thickness. Alternatively, the article
or component may have mechanical properties that vary across the
article by providing distinct regions, each region having different
mechanical properties than an adjacent region by providing ceramic
fibers in one region having a different coating thickness than
ceramic fibers in an adjacent region. In this embodiment, the
mechanical properties within any one region may be substantially
uniform.
In a further embodiment, the present invention provides a ceramic
matrix composite with tailored properties that includes a ceramic
matrix, a plurality of ceramic fibers disposed through the ceramic
matrix, wherein the matrix is bonded to the fibers, and a coating
on the ceramic fibers, wherein the coating thickness is
instrumental in determining the strength of the bond between the
fibers and the matrix, the coating thickness being varied in a
predetermined manner on adjacent regions of fibers to produce
adjacent regions of varying strength. The coating thickness may
approach or be zero and may be a monolayer of a single coating
material having a thickness of about 1 micron (.mu.). The coating
may also be applied as multiple layers of coating material of
different compositions to a thickness of about 3.mu., each layer
having a thickness of up to about 1.mu..
Other features and advantages of the present invention will be
apparent from the following more detailed description of the
preferred embodiment, taken in conjunction with the accompanying
drawings which illustrate, by way of example, the principles of the
invention.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a perspective view of a turbine blade of an aircraft
engine.
FIG. 2 is an enlarged sectional view of blade of FIG. 1, taken
along the line 2-2.
FIG. 3 is a calculated ANSYS.TM. interlaminar stress pattern of the
dovetail portion of the blade of FIG. 1 taken along line 3-3,
indicating the different stress regions across the dovetail
cross-section.
FIG. 4 is a graphical representation of the fiber coating thickness
as measured along line 4-4 of FIG. 3.
FIG. 5 is a flow chart illustrative of a method in accordance with
the present invention.
DETAILED DESCRIPTION OF THE INVENTION
FIG. 1 illustrates a turbine blade 10 such as typically used in a
gas turbine engine (not shown). The turbine blade includes an
airfoil portion 12 and a dovetail portion 14. Although a turbine
blade may optionally include a platform portion positioned between
the airfoil portion 12 and the dovetail portion 14, the embodiment
shown in FIG. 1 does not include the optional platform portion.
Airfoil portion 12 is defined by the airfoil, which extends from
the airfoil tip 16 toward an intermediate portion 18, where airfoil
portion 12 widens into dovetail portion 14. Dovetail portion 14
includes a contoured outer surface 20 that is used to secure the
turbine blade 10 in the disk or rotor (not shown) of the gas
turbine engine. The dovetail portion seats into a corresponding
dovetail slot formed in the periphery of the rotor. It will be
appreciated that blade 10 experiences maximum mechanical stresses
within dovetail portion 14 during engine operation and that thermal
stresses within dovetail portion 14 are greater in the hotter
sections of the engine as the blade dovetail is pushed against the
interfacing forces of the disk by the rotational forces of the
engine. It will be appreciated that the maximum rotationally speed
of the turbines occurs when the temperatures produced by the
combustor are highest, as more fuel is burned to produce more
power. In accordance with this invention, blade 10 is a ceramic
matrix composite constructed of ceramic fibers 22 provided with an
interface coating 24, the coating applied to the fibers to form an
interface between the fibers 22 and a matrix 26, infused into the
interstitial spaces between fibers 22.
FIG. 2 is an enlarged view of the internal structure of the CMC
blade 10, illustrating, in simplistic form, ceramic fibers 22 with
an interface coating 24 applied thereon. Preferably, the ceramic
fibers 22 are SiC. However, the present invention is not limited to
turbine blade structures, and other uses of the present invention
both within a gas turbine engine and in other applications are
envisioned. Another application within a gas turbine engine is, for
example, as a compressor blade in the compressor, which operates at
considerably lower temperatures. Other suitable fibers that may be
used, depending on the specific application, include silicon
nitride, aluminum oxide, silicon-nitrogen-carbon, silicon carbide
sheath overlying a carbon core, aluminum borate, silicon oxide,
silicon carbide that includes a metal, such as titanium or nickel,
silicon oxycarbides, carbon and the like.
With reference to FIG. 3, dovetail portion 14 is illustrated in
greater detail. FIG. 3 further illustrates the calculated ANSYS.TM.
interlaminar stresses within nine separate regions of dovetail
portion 14. These regions, labeled 30, 32, 34, 36, 38, 40, 42, 44,
and 46, experience differing interlaminar and in-plane stresses
during engine operation. In the example provided, region 30
experiences high interlaminar stress and low in-plane stress when
compared to the other regions. Similarly, region 46 experiences
high in-plane stress and low interlaminar stress when compared to
the other regions. The interlaminar stresses decrease progressively
from region 30 to region 46, while the in-plane stresses increase
progressively from region 30 to region 46. It is clear from FIG. 3,
that different portions of the article, here a turbine blade, are
subjected to different stresses, and materials properties that are
suitable for use in one location may be unsuitable in another
location under the same environmental conditions.
FIG. 4 graphically illustrates the thickness of interface coating
24 on fibers 22, measured along the line 4-4 of FIG. 3 for an
embodiment of the invention. Line 50 graphically represents the
thickness of coating 24 in each region 30, 32, 34, 36, 38, 40, 42,
44, and 46, wherein region 30 has a very thin application of
interface coating 24 than other regions, the interface coating
increasing with in-plane stresses inversely to interlaminar
stresses. The thickness of the interface coatings 24 varies from a
minimum or zero in region 30 where the in-plane stresses are at a
minimum. The maximum thickness of the interface coating can vary
dependent upon the make-up and form of the interface coating. When
the interface coating comprises a plurality of thin layers of
differing composition, the thickness of the layers can be up to 3.0
microns (.mu.). When the interface coating is a monolithic layer of
BN, the thickness of the coating typically is 1.mu.. The thickness
of coating 24 will depend upon the types of stresses and the value
of the stresses present in the region in which the fiber is
located. Thus the range of thickness of the coating is stress
dependent, and will vary from application to application. For an
article having a stress distribution pattern such as the blade
cross-section depicted in FIG. 3, the thickness of the coating
varies in a manner as shown in FIG. 4.
For an article such as a turbine blade, having a coating thickness
with a thickness pattern such as shown in FIG. 4, the coating
distribution pattern can be achieved in any acceptable manner. In
one embodiment, the coating distribution of the fibers can readily
be achieved by drawing the fibers through a solution of the
coating. Thickness of the coating can be varied by controlling the
dwell time of the fibers in the solution. A thicker coating is
achieved by varying the speed at which the fibers are drawn through
the solution. Once the fiber tows have been coated in this manner,
the fibers can be drawn through a solution of matrix to form a
prepreg sheet or ply. The prepreg plies are then cut and laid up so
that the prepreg plies having a thin coating are located in the
region of the layup having the highest interlaminar stresses, while
the prepreg plies having the thickest coating are located in the
region having the highest in-plane stresses. After the desired
lay-up is achieved, the prepreg plies are consolidated to form a
ceramic matrix composite by application of heat and pressure for a
preselected time. Consolidation is preferably accomplished by
applying a pressure of about 100-250 psi at a temperature of about
100-200.degree. C. (about 212-392.degree. F.) for about 12-36
hours. The final ceramic matrix composite blade that results from
this process is a blade in which there is some bonding between the
matrix and the fibers in the areas in which the interlaminar
stresses are high. The material acts monolithically in this region,
which is desirable in the region of interlaminar stresses. In areas
where the in-plane stresses are high, there is a thick coating
between the matrix and the fibers, which may result in some
slippage between the matrix and the fibers. The coating allows the
load to be transferred from the matrix to the fibers so that the
stress can be transferred to the fibers so as not to overstress the
matrix, which could fail in a brittle manner.
In another embodiment, the distribution of coating thickness of the
fibers in which the coating thickness of the inner fibers, or the
fibers comprising the inner portion of the preform, is thinner than
the coating thickness of the outer fibers, or fibers comprising the
outer portion of the preform, such as is required for a turbine
blade, can be achieved as follows. A fiber preform using uncoated
fibers is first laid up. The fiber preform is then exposed to a
chemical vapor infiltration process in which the vapor comprises
the coating material. The coating vapor infiltrate the preform,
depositing coating material on the unexposed fibers. The thickness
of the deposited coating will be thinner in the inner fibers of the
preform and thicker on the exterior fibers of the preform. The
coating thickness can be further varied by varying the partial
pressure of the coating vapor so as to control the concentration of
vapor reaching the interior fibers. After the desired coating
thickness and distribution has been achieved, the preform can be
melt-infiltrated with matrix material using the slurry cast melt
infiltration process so as to achieve a ceramic matrix composite
material.
This technique is effective for an article such as a turbine blade
in which the coating thickness of the fibers on the interior of the
preform is thinner than on the exterior. However, when the stress
distribution pattern of the article is reversed, requiring a
thicker coating applied to the fibers on the interior of the
article preform, the coating thickness distribution can be achieved
by applying a mask to fibers on the exterior of the preform before
exposing the preform to the coating vapors. After the desired
coating thickness has been achieved in the interior of the article
preform, the mask material can be removed and the fibers on the
exterior of the preform can be coated. Although the thickness
distribution will not be a mirror image of the distribution shown
in FIG. 4, the general pattern will provide a thicker coating in
the interior and a thinner coating in the exterior of the preform.
The preform can then be slurry cast melt infiltrated in the
conventional manner.
In yet another embodiment, fibers can be provided with a coating
and a preimpregnated (prepreg) matrix, the fibers grouped in
accordance with the coating thickness applied to the fibers. The
coated fibers in each grouping can have a uniform coating thickness
along their length, or the fibers may have a coating thickness that
varies along the fiber length. This variable thickness can be
obtained, for example, in a manner such as described above for
formation of prepreg sheets. The article can be subdivided into a
plurality of sectors, such as the eight sectors represented by
regions 30-46 in FIG. 3. The fibers can be grouped into an
appropriate number of sectors, here eight sectors, according to the
coating thickness applied to the fibers. Each group of fibers can
then be assembled into a preform so that a plurality of fibers
having a thin interface coating, or no interface coating, is
assembled at a location corresponding to region 30. Then,
additional groups of fiber are assembled into regions 32-46 as a
function of coating thickness, with the fibers with the thickest
coating assembled in region 46, thereby providing a preform having
fibers with a stepwise coating thickness gradient. It will be
appreciated that while this technique provides a gradient that
varies from a thin (or no) coating in the interior, to a thick
coating on the exterior corresponding to low interlaminar stresses
on the exterior and high interlaminar stresses in the interior, the
fibers can be arranged to correspond to high interlaminar stresses
on the exterior and low interlaminar stresses in the interior, or,
if dictated by the stress distribution pattern, alternating high
and low interlaminar stresses across a cross-section. It will also
be appreciated that the coating thickness along the fiber length
can be graded, so that one portion of a fiber or group of fibers
may be in a section of high interlaminar stresses, while another
portion of a fiber or group of fibers may be in a section of low
interlaminar stresses. Thus, the preform layup can reflect the
variation in thickness along the length of the fibers (and hence
the axis of the article) as well as across the cross section of the
part. After the preform has been assembled with the prepreg sheets,
the preform can be melt infiltrated to achieve a ceramic matrix
composite material as previously discussed.
Preferred diameters for the fibers used in the composites of the
present invention vary in the preferred embodiment from about 5-20
microns (.mu.) in diameter, and most preferably from about
10-15.mu. in diameter. These values do not include coating
thicknesses, the application of which may vary from location to
location within the CMC article. The applied coating thickness (t)
will increase the diameter of the fiber by the value of 2t,
depending on the coating thickness at a given location. Since in
certain locations the coating thickness may be as high as 5.mu.,
the overall diameter of the fibers will be accordingly increased by
as much as 2t or 10.mu., or from 10.mu. to 25.mu..
As noted above, the fiber coatings may be monolithic BN. The fiber
coating thickness will vary depending upon the stress intensity and
type of stress (in-plane v. interlaminar) to which the fiber is
subjected. For a monolithic BN interface coating, the interface
coating thickness can vary from no coating in a region of high
interlaminar stresses (and low in-plane stresses to a coating
thickness of about 1.mu. in regions subjected to in-plane
stresses.
It would be appreciated that the process described herein, while
preferably directed to SiC/SiC composites with a BN fiber coating,
may be accomplished with other composite and coating materials to
produce the desirable material properties. For example, because
certain processing of the preform into a ceramic matrix composite,
such as for example infiltration processes, can cause damage or
deterioration to BN coatings, it is sometimes desirable to apply
the interface coating as a plurality of layers. In one embodiment,
the coating comprises four layers. Each layer has a specific
purpose, and fewer or more layers may be used as required. Of
course, such an interface coating is desirable in regions in which
it is important to maintain the BN coating as a distinct coating.
Thus, the interface coating applied as a plurality of layers is
preferred in regions in which there is high in-plane stress,
because it is desirable to prevent the matrix from interacting with
the fibers, thereby allowing the stresses to be transferred from
the matrix to the fibers. The layer immediately adjacent to the
fibers is BN, which is useful in providing the slip between the
fiber and the matrix. However, BN is susceptible to oxidation,
particularly in the presence of H.sub.2O (steam), thus a more
oxidation resistant Si--BN overlayer is applied. The BN-containing
layers have a tendency to react with molten silicon carbide. Thus,
to protect the BN coating and Si-doped BN overlayer, an overlayer
of silicon nitride (SiN). Is applied. Optionally, a layer of carbon
is applied over the SiN. The optional carbon layer, applied to a
thickness up to about 0.1.mu., and preferably about 0.05.mu.
promotes interaction at the interface of the SiN with any molten
Si. The carbon promotes infiltration of the molten silicon,
improving preform "wetting" with silicon and formation of SiC. The
SiN may decompose if molten Si penetrates it, but again this
decomposition is designed to protect the underlying BN. This layer
is applied to a thickness of about 0.75.mu.-1.25.mu.. The Si doped
BN layer is applied to a thickness in the range of about
0.4-0.6.mu.. In areas in which there are high interlaminar
stresses, and interactions between the fibers and the matrix are
helpful in preventing the fibers from acting as defects with
respect to the stresses, such multilayer coatings are not desired,
as they can inhibit such interactions. All of the fibers in a
preform may be coated with the multilayer system or only
preselected fibers may be coated with the multilayer system. The
coating systems can thus be varied depending upon the stress
distribution patterns that are anticipated and the interactions
required between the fiber and the matrix as a result of these
stresses.
The overall coating thickness of the above-described layer can
approach 3.mu., which is effective in promoting a maximum transfer
of stress to the fiber. Less transfer can be achieved with a
thinner coating. For an intermediate system, in which stress is
transferred to the fiber, but not the maximum amount possible, BN
can be applied to a thickness of about 0.3.mu., the Si-doped layer
can be applied to a thickness of about 0.3.mu. and the SiN can be
applied to a thickness of about 0.5.mu.. Of course, minimal
transfer of stress can be obtained by allowing the SiC fiber to
react with the SiC matrix, even to the point of forming a
monolithic material by allowing a complete reaction.
The acceptable coating are not limited to the BN or to the
multilayer system described above. Other acceptable coatings that
can be substituted entirely or combined in layers with BN layers as
part of a multi-layer system include for example, Si.sub.3N.sub.4,
SiC, aluminum nitride and carbon. These coatings can be applied
individually up to 1.mu. or as coatings in a system wherein the
thickness of the coating system can be up to 3.mu..
While SiC is a preferred matrix, other matrix materials may also be
used with the graded coated fibers of the present invention. Two
other acceptable matrices include silicon-nitrogen-carbon matrices
and aluminum silicates matrices. Articles can be fabricated from
these matrices using the above slurry cast melt infiltration
techniques, the prepreg techniques described above, or any other
acceptable technique, such as polymer impregnation pyrolysis (PIP).
Various organosilicon preceramic polymers may be used to form the
silicon based matrices. These prepolymers include, but are not
limited to polysiloxanes, polysilazanes, polysilanes,
polymetallosiloxanes and the like. The material used to form the
matrix is not important, as long as the matrix can be formed around
the fiber to achieve the desired result depending upon the type and
intensity of stresses anticipated and experienced at the interface
between the fiber and the matrix, as discussed above.
Referring now to FIG. 5, there is shown a flow chart illustrating a
slurry cast MI method of manufacture an embodiment of the present
invention to produce a CMC turbine blade. The initial step 100 of
the process is laying up a preselected number of biased SiC
containing cloth plies of preselected geometry in a preselected
arrangement to form a turbine blade shape, or preform. In a
preferred embodiment, there are a preselected number of fiber tows
woven in the weft direction sufficient to allow the SiC cloth to be
handled and laid up without falling apart. A CMC element
manufactured with biased SiC containing cloth plies has greater
tensile strength in the warp direction of the SiC containing cloth
plies than the weft direction. The tensile strength in the warp
direction is up to about 25 percent greater than in the weft
direction.
Once the plies are laid up, the next step 110 is to rigidizing the
turbine blade shape with BN. As is known, an applied BN coating on
SiC fibers in a SiC matrix provides a weaker bond between fibers 22
and matrix 26. In step 110, the thickness of coating 24 is varied
across differing regions, such as regions 30, 32, 34, 36, 38, 40,
42, 44 and 46 of FIG. 3, to tailor the strength within each region
to desired amounts. The desired amounts of coating required for
fiber in a designated region may be known from extensive
destructive testing, or from computer simulations that predict the
stresses, and therefore desired strengths, within each region. The
coating thickness can be varied by any of the techniques set forth
above that will produce a uniform and predictable thickness for the
fiber in the region. In a preferred embodiment, the method includes
applying BN coatings using a chemical vapor infiltration (CVI)
process, forming a rigid coated turbine blade preform. Thus
provided, the thickness of BN can be varied across the preform
regions 30, 32, 34, 36, 38, 40, 42, 44 and 46 to increase desired
strengths in desired regions. A uniform coating thickness of BN
applied to the fibers in a preform is known to provide the final
CMC component with improved mechanical properties, including
improved modulus of elasticity, improved tensile strength, and
improved fracture toughness through the component, however, very
few components experience consistent levels of stresses throughout.
However, a preform having fibers with graded thicknesses should
better be able to survive the varied types of stresses (i.e.
interlaminar, in-plane through-thickness) experienced by the
article or component in service as compared to fibers having a
uniform coating thickness.
The next step 120 partially densifies the coated turbine blade
preform by introducing a carbon-containing slurry, as is known in
the art, into the void areas between the fibers of the coated
turbine blade preform. The final step 130 further densifies the
turbine blade preform with at least silicon, and preferably boron
doped silicon, through a slurry cast MI process, in which the
silicon reacts with carbon to form a SiC matrix, the final part
being a SiC/SiC CMC turbine blade with biased architecture.
For a preselected fiber/matrix combination, the grading of the
coating will be varied depending upon the measured or predicted
stress patterns and intensities to permit the use of CMC's in
applications in which they were heretofore not able to be used.
Thus, the grading of the coating applied to the fibers to control
the interaction of the fibers with the matrix based on measured or
calculated stress patterns and intensities will vary depending upon
the location, and therefore the stress pattern within the CMC
article. For example, a thin coating may be applied in an area of
the article where stress intensity for interlaminar stresses is
high, while an adjacent region may have a thick coating to
accommodate higher in-plane stresses. So therefore, coating
thickness may vary from as low as 0.5.mu. total to as high as 5.mu.
total.
While the invention has been described with reference to a
preferred embodiment, it will be understood by those skilled in the
art that various changes may be made and equivalents may be
substituted for elements thereof without departing from the scope
of the invention. In addition, many modifications may be made to
adapt a particular situation or material to the teachings of the
invention without departing from the essential scope thereof.
Therefore, it is intended that the invention not be limited to the
particular embodiment disclosed as the best mode contemplated for
carrying out this invention, but that the invention will include
all embodiments falling within the scope of the appended
claims.
* * * * *